MULTI-DISCIPLINARY DESIGN INVESTIGATION OF PROPULSIVE FUSELAGE AIRCRAFT CONCEPTS

Transcription

1 MULTI-DISCIPLINARY DESIGN INVESTIGATION OF PROPULSIVE FUSELAGE AIRCRAFT CONCEPTS JULIAN BIJEWITZ, ARNE SEITZ and MIRKO HORNUNG Bauhaus Luftfahrt e.v., Lyonel-Feininger-Straße Munich, Germany Abstract Motivated by the potential of gaining noticeable improvements in vehicular efficiency, the benefits attainable from introducing a more synergistic integration of the propulsion system to the airframe is investigated in this paper. In previous work, the concept of a boundary layer ingesting propulsor encircling the aft section of an axisymmetric fuselage was identified to be particularly promising for the realization of aircraft wake filling, and hence, a significant reduction of the propulsive power required. After reviewing the theoretical foundation of the Propulsive Fuselage concept, a book-keeping and model matching procedure was introduced, which was subsequently used to incorporate the numerical aerodynamic characteristics of a Propulsive Fuselage aircraft configuration into a propulsion system sizing and performance model. As part of this, design heuristics for important characteristics intrinsic to Propulsive Fuselage power plants are derived. Thereafter, parametric study results of the propulsion system are discussed and the obtained characteristics are compared to those of a conventionally installed power plant. Finally, the impact of the investigated propulsion system on the integrated performance of a Propulsive Fuselage aircraft concept is studied, and the results are compared and contrasted to previously conducted analyses based on semi-empirical characteristics. Keywords Distributed Propulsion, Propulsive Fuselage, Boundary Layer Ingestion, Aircraft Wake Filling, Conceptual Aircraft Design 1. Introduction In order to close the gap between the ambitious long-term environmental targets outlined by the European Commission (EC, 2011; ACARE, 2012) and the improvements attainable from incremental enhancement of conventional technology, an exploration of the efficiency potentials and feasibility of novel options for propulsion system design and synergistic aircraft integration is highly warranted. Particular aircraft-level benefits are expected from the prospect of distributing the production of thrust along main components of the airframe, i.e. distributed propulsion (Kim, 2010). A most promising concept for distributed propulsion is the Propulsive Fuselage (PF) concept (Steiner et al., 2012), which is currently subject to a multi-disciplinary investigation in the EU-funded Level 0 project DisPURSAL (Isikveren, 2012; Isikveren et al., 2014). Key element of the concept is a single large propulsor encircling the aft part of a cylindrical fuselage with intent to ingest the fuselage boundary layer, thereby allowing for aircraft wake filling to the greatest extent. A number of variations of the general PF idea can be found in the literature (Bolonkin, 1999; Stückl, 2012; Schwarze, 2013). Seitz and Gologan (2013, 2014) introduced a unified book-keeping scheme of system-level efficiency figures applicable to highly integrated boundary layer ingesting propulsion systems and conducted an initial sizing study based on semi-empirical modelling of the fuselage boundary layer ingestion (BLI). Kaiser et al. (2014) introduced a set of quasi-analytical aerodynamic methods for efficient exploration of PF configurations. Seitz et al. (2014) focused on the conceptualization and preliminary sizing of a PF aircraft layout and identified a best and balanced design yielding an increase in vehicular efficiency of approximately 10% compared to an advanced reference aircraft targeting year 2035 Entry-into-Service (EIS). In the present paper, a procedure for the incorporation of numerical aerodynamic characteristics of a PF configuration into a propulsion system design model is introduced. Based hereon, parametric design studies of a 1

2 2 J. Bijewitz, A. Seitz and M. Hornung PF power plant are presented and discussed. Finally, the suitability of the method for analyses at integrated vehicular level is demonstrated. 2. Theoretical Basis This section provides a brief overview of the general implementation of the PF concept and the associated propulsion system. Important metrics for performance assessment of boundary layer ingesting and conventional power plants, as well as evaluation of the vehicular efficiency are reviewed The Propulsive Fuselage concept In Figure 1, a PF integration concept as selected in (Seitz et al., 2014; Isikveren et al., 2014) is presented. The configuration is based on a three-engine layout. The aft-installed power plant is primarily intended to serve the purpose of wake filling, while conventionally installed turbofans deliver the residual thrust required to operate the aircraft. By ingesting the fuselage boundary layer and re-energizing the momentum deficit in the aircraft viscous wake caused by skin friction on the wetted fuselage surface, a reduction of propulsive power can be attained (Seitz and Gologan, 2014). In the investigated concept, the so-called Fuselage Fan, FF, installed at the aft of the fuselage is driven by a turbo engine installed in the fuselage aft cone, which is supplied with air through an s-shaped inlet duct installed downstream of the FF (see Figure 1, right). The FF is powered by the Low Pressure Turbine, LPT, via a planetary reduction gear system. Further key aspects of the propulsion system concept, in particular regarding integration aspects, are discussed in detail by Seitz et al. (2014). In view of a potential Entry-Into-Service, EIS, year 2035, technology freeze was set as Fuselage Fan Gas turbine installation position z Intake Strut Fuselage Fan Rotor Fan Stator Vertical Tail S-duct (Core Intake) Core Exhaust x Fuselage Fan Drive Gear System Booster High Pressure Compressor Low Pressure Turbine High Pressure Turbine Figure 1: Propulsive Fuselage aircraft concept (left, adopted from Isikveren et al., 2014) and Fuselage Fan propulsion system (right) 2.2. Review of metrics for performance evaluation For consistent treatment of conventionally installed, i.e. under wing podded, and highly integrated propulsion systems such as the PF concept, the definition of unified efficiency standards is required. Seitz and Gologan (2014) introduced a scheme for the definition of propulsion system efficiency figures applicable to conventional and FF power plants. Here, the interface of thrust/drag book-keeping between the propulsion system and the airframe is geared to the propulsion system streamtube. Hence, all aerodynamic effects in the streamtube ahead of the FF inlet are incorporated in the power plant sizing and performance analysis. As will be shown later in the paper, the breakdown of overall propulsion system efficiency, η ov, into the three individual contributors, i.e. the core efficiency, η co, the transmission efficiency, η tr, and the propulsive efficiency, η pr, is convenient in order to identify the impact of BLI on power plant performance: PThrust FN V0 ov co tr pr (1) P P Supply Supply where F N denotes the net thrust and V 0 the free stream velocity. A detailed discussion of the individual efficiencies in the context of the PF concept is provided by Seitz and Gologan (2013, 2014). Nacelle external aerodynamics are considered to belong to the aircraft characteristics, thus influencing the thrust required to operate the aircraft. The amount of aircraft drag captured inside the propulsion system streamtube and hence

3 Multi-Disciplinary Design Investigation of Propulsive Fuselage Aircraft Concepts 3 taken as being removed from the aircraft drag balance, D ing, is described using the ingested drag ratio (cf. Steiner et al., 2012), Ding (2) F N, t where F N,t represents the total net thrust required for aircraft operation. The assessment of vehicular performance may be performed using the Energy Specific Air Range, ESAR, indicating the change of aircraft range, R, per change of energy, E (cf. Seitz et al., 2012): dr V0 L / D ov L / D ESAR de TSPC m g m g A / C including the aircraft aerodynamic efficiency, L/D, its mass, m A/C, and the propulsion system overall efficiency. A / C 3. Book-keeping and Matching Procedure for Propulsive Fuselage Concept The present section firstly summarizes the setup used for the numerical calculations of the PF concept. Thereafter, the method used for modeling of the FF propulsion system is discussed and a procedure for the matching of CFD results with gas turbine performance calculations is introduced Description of the CFD setup An essential task associated with a PF concept refers to the aerodynamic assessment of the airframe-propulsion integration. During the research activities associated with the DisPURSAL Project, numerical flow computations of the PF configuration were performed by the French research institute Office National d'études et de Recherches Aérospatiales, ONERA (Isikveren et al., 2014). An overview of the institute s expertise in this field including the validation of the aero-numerical methods applied in the present context can be found in (Atinault et al., 2013). Pursuing the purpose of an initial design space exploration, two-dimensional RANS calculations of an axisymmetric arrangement comprising the fuselage and the FF nacelle were conducted for a representative cruise condition (FL350, M0.80). Hence, the flow received into the FF inlet is assumed to be orthogonal to the fan vertical plane. The pressure increase due to the FF power plant was simulated using an appropriately adapted actuator disk model. Potential detrimental interference emanating from the vertical tail attachment disregarded in the CFD analysis was accounted for in the aircraft sizing studies using an appropriately selected interference factor. A detailed description of the computational setup, imposed assumptions and main results is given in (Isikveren et al., 2014). Aerodynamic computations were conducted for a number of PF power plant designs based on an experimental plan designed by Bauhaus Luftfahrt. The correspondingly investigated PF geometries were generated by Airbus Group Innovations, AGI. For each geometric configuration, a number of FF rotational speeds were simulated in order to examine different power settings. Important results relevant for aircraft and propulsion system sizing such as performance parameters, flow Mach numbers and pressure ratios were fed back to Bauhaus Luftfahrt for further data analysis and subsequent system sizing. As part of the CFD data postprocessing, a common design axial fan inlet Mach number, M ax,2,des, was identified for the sampled PF designs. M ax,2,des typically constitutes an important similarity parameter for gas turbine sizing and performance mapping. Based on an interpolation of the mass flow averaged fan inlet Mach numbers of the simulated PF designs at different power settings, M ax,2,des = 0.56 was identified suitable in order to ensure PF design similarity. This fan inlet Mach number is approximately 20% below the value typical for conventionally installed advanced turbofan propulsion systems (Grieb, 2004) Modeling of the propulsion system For propulsion system sizing and performance simulation the gas turbine performance software GasTurb 11 (Kurzke, 2010) was utilized. For cycle and flow path sizing typical design laws based on Seitz (2012) and Grieb (2004) were applied including iteration of design net thrust, Overall Pressure Ratio (OPR), outer and inner Fan Pressure Ratios (FPR), fan tip speed as well as nozzle gross thrust and discharge coefficients. Consistent settings (3)

4 4 J. Bijewitz, A. Seitz and M. Hornung for the parametric mapping of High Pressure Turbine (HPT) cooling air demand as a function of hot gas, bulk material and cooling air temperature were implemented using the approach given by Seitz (2012). Turbo component efficiencies, basic cycle characteristics as well as duct pressure losses were adjusted in order to reflect an EIS of year Technology settings were retained constant throughout the studies performed. A best and balanced OPR was determined with respect to Turbine Entry Temperature (TET) at take-off and corresponding cooling air demand. Compressor work split was chosen to allow for uncooled LPTs. In the first instance, core size effects on the efficiency of turbo components were neglected. Gearbox losses intrinsic to a Geared Turbofan (GTF) architecture were incorporated in the mechanical losses of the low pressure spool. Propulsion system dimensions and weight estimation referred to a component build-up approach described in detail in (Seitz, 2012). The estimation of propulsor drive gearbox masses was based on Reynolds (1985), however, calibrated to reflect advanced technology status. Since the FF features a very high hub-to-tip ratio compared to podded fan designs, the weight of the FF module was estimated using the method described in (Steiner et al., 2012). For the fans of both the podded power plants and the FF propulsion system a material mix consisting of 20% titanium and 80% CFRP was chosen corresponding to advanced technology standard (Seitz et al., 2014). For the modelling of the FF propulsion system, a number of adjustments relative to the reference power plant were required. The fan hub-to-tip ratio was iterated according to the duct height and the prescribed hub radius, which was assumed to equal the fuselage radius at the given axial position. Due to the high hub-to-tip ratio of the FF compared to conventionally installed power plants, outer and inner fan pressure ratios were assumed to be identical Matching of CFD and system performance prediction The following sections initially provide an overview of the design space investigated in the present context. Furthermore, the approach for the mapping of key characteristics associated with the integration of the FF propulsion system is discussed Investigated design space In order to maximize computational throughput, emphasis was placed on investigating the impact of the most important free variables applicable to a FF power plant design. Apart from the FF inlet duct height, which has been identified as an essential design variable (cf. Seitz and Gologan, 2014; Seitz et al., 2014), FPR was considered important, since it governs the level of specific thrust produced by the power plant and is, therefore directly related to η pr. While for given duct heights specific values were prescribed explicitly for the aeronumerical analyses, FPR was set implicitly by specifying the required nozzle exit area obtained from initial GasTurb calculations. As a first step, an initial geometry was optimized in a multi-disciplinary design effort according to the workflow described in detail by Isikveren et al. (2014). Here, particular emphasis was placed on the identification of an optimum FF nacelle shape for minimum intake spillage drag, and, the optimization of the fuselage body contour in front of the FF in order to avoide unfavorable flow conditions such as localized supervelocities or flow separation. Using this optimized configuration, a design space exploration of the described parameters was performed. The feasible range of duct heights at the Aerodynamic Interface Plane (AIP), h AIP, was chosen from 0.50 m to 0.90 m. Due to the employed actuator disc model, FPR was limited to For the lower feasible bound, a value of 1.20 was selected. During the investigations, the distribution of design points was continuously subject to adaptation based on the results obtained for each previous design point. Fuselage shaping was retained as defined during the initial optimization (denoted as D 0 ). Also, the general nacelle parameterization was kept invariant, in the first instance, and a constant offset between the tip radii at the AIP, and the first actuator disk was implemented. Based hereon, five additional design points (D 1 through D 5 ) were evaluated through CFD and used for the subsequent derivation of propulsion system (PPS) sizing and performance heuristics. The computed designs were distributed in the design space as given in Figure 2.

5 Multi-Disciplinary Design Investigation of Propulsive Fuselage Aircraft Concepts 5 Duct Height at AIP (h AIP ) [m] D 2 D 3 D 1 D 4 D 5 D Design Fan Pressure Ratio (FPR) [-] Figure 2: Investigated design space Matching procedure Exploration of the PF design across the entire feasible design space requires a propulsion system model capable of reflecting the system behavior within the corresponding range of design parameters. In order to enable continuous power plant calculations over a wide range of design parameters, the results derived from the aeronumerical experimentation described above and the subsequent post-processing procedure were used for adaptation of the PF sizing and performance model simulated in GasTurb. The scope of the GasTurb model was tailored to cover all streamtube effects for the FF power plant system (cf. Seitz and Gologan, 2014). Now, the target of the matching procedure was to incorporate the physical effects derived from the aero-numerical analysis, such that the response of the GasTurb model is consistent with the CFD results obtained for the sample points (D 0, D 1 through D 5 ). Therefore, regression models were derived for important parameters affected by the present application, such as the intake pressure ratio (p t2 /p t0 ). Subsequently, these heuristics were integrated to the propulsion system design and performance model which was finally wrapped for aircraftintegrated simulation using surrogate modelling techniques. The employed workflow is schematically presented in Figure 3. Aero-numerical Data (Isikveren et al., 2014) Nonlinear Regression of derived parameters (MATLAB ) Semi-empirical Boundary Layer Method (Based on Seitz and Gologan, 2014) Verification PPS Design Model matched to Aeronumerical data (GasTurb 11) Parametric PPS Design Studies (GasTurb 11) PPS Surrogate Modelling (Seitz, 2012) Figure 3: Procedure used for CFD and gas turbine performance matching Modelling of intake pressure ratio As a consequence of the momentum deficit formed by fuselage skin friction in front of the intake of a PF arrangement, the total pressure at the intake is decreased relative to the total pressure at undisturbed free stream velocity. Hence, the intake total pressure ratio is degraded compared to a conventionally installed power plant. Depending on the specific geometry investigated, the impact on the pressure ratio can be significant (Seitz and Gologan, 2014).

6 6 J. Bijewitz, A. Seitz and M. Hornung As h AIP and FPR constituted the variables investigated in the aero-numerical investigation, the influence of these parameters on p t2 /p t0 was investigated. As recognized in a previous investigation (Seitz and Gologan, 2014), p t2 /p t0 was found to be highly sensitive with h AIP. In contrast to that, the dependence on design FPR was identified to be small. Hence, only the sensitivity of p t2 /p t0 with duct height was included in the approximation model, which was based on a polynomial approach and is presented in Figure 4. The maximum approximation error was 0.17%, while the mean error was 0.12%. The model is considered valid for duct heights between 0.3 m and 1.1 m. The derived correlation is included in the figure Data points obtained from CFD analysis 2 Polynomial Approx.: p t2 /p t0 = *h AIP *h AIP, valid in haip = [0.5m, 0.9m] Intake Pressure Ratio (p t2 /p t0 ) [-] D 5 D 4 D 0 D 1 Study Settings: Operating Condition: FL350, M0.80, ISA Fuselage Geometry acc. to Isikveren et al. (2014) Rel. Longitudinal Position of AIP: 85% Fuselage Length: 69 m Fuselage Equivalent Diameter: 6.07 m D 2 D Duct Height at AIP (h AIP ) [m] Figure 4: Implemented approximation model for intake pressure ratio The sensitivity of flight Mach number was only analyzed for the investigation associated with the initially optimized geometry (D 0 ). In order to still allow for a parametric mapping of flight Mach number, the sensitivity obtained from the D 0 design was considered representative, in the first instance. From that, normalized scaling factors were determined and introduced into the model given in Figure 4. The obtained trend causes the pressure recovery to decrease slightly for increasing flight Mach numbers, M 0. This is intuitive as the intake total pressure ratio of a boundary layer ingesting power plant is dominated by the ratio of the mean total pressure at the intake, p t1, and the respective value at free stream condition, p t0. This may be expressed through an equivalent isentropic ram pressure recovery factor (cf. Seitz and Gologan, 2014): pt p t M 2 2 1, m M where M 1,m represents the mean intake Mach number. Despite the fact that the boundary layer thickness decreases with increasing free stream Mach numbers, and hence p t1 tends to increase, the effect is overcompensated by the strong correlation between p t0 and M Mapping of integration losses An important implication associated with a PF configuration refers to the additional drag effects caused by the presence of the FF power plant at the aft fuselage. The CFD-based design space exploration had shown that FF intake spillage drag is highly sensitive to both fuselage and nacelle contour definition. Beyond that, the FF nozzle shear flow across the fuselage aft-cone creates additional drag. As the setup for the aero-numerical computation did not facilitate a component based analysis of the individual drag effects, a conversion procedure (4)

7 Multi-Disciplinary Design Investigation of Propulsive Fuselage Aircraft Concepts 7 was employed in order to identify the individual contributors to the overall thrust/drag balance. The matching procedure applied for a consistent transformation of thrust values and the quantification of integration drag effects is briefly outlined below. Therefore, in Figure 5 a scheme of a generic PF layout is depicted indicating the propulsion system streamtube, the fuselage boundary layer, as well as forces and corresponding control volumes, thus illustrating the thrust/drag book-keeping used in the present context. 0 AIP Control Volume Boundary Layer Velocity Profile Fuselage Body D jsf Propulsion System Streamtube D Fus, ing D Fus,not ing Fuselage Boundary Layer D Nac Fuselage Drag Ingested (D Fus,ing ) Fuselage Drag not Ingested (D Fus, not ing ) Drag due to Jet Shear Flow (D jsf ) Thermodynamic Station Numbering D Nac Thrust of Propulsor Nacelle Drag Figure 5: Scheme used for thrust/drag book-keeping The definition of net thrust measured in the frame of reference used for the CFD simulation, Fˆ N, is given by the thrust produced by the actuator disks, Fˆ, reduced by the drag of the configuration accounted for in the CFD: ˆ F N Fˆ Dˆ (5) where variables denoted with the hat symbol indicate parameters referring to CFD calculations. In the context of the CFD setup, the fuselage drag, D Fus, the nacelle drag, D Nac, as well as jet shear flow losses occurring at the fuselage aft-cone, D jsf, constitute the contributors to Dˆ. Thus, equation (5) yields: F ˆ F ˆ D D D (6) N Nac Fus Depending on the boundary layer thickness and the duct height of the propulsor, not necessarily the entire fuselage viscous drag is ingested into the propulsive device. Hence, the fuselage drag may be split up into: jsf D Fus DFus, ing DFus, noting (7) where D Fus,ing represents the fuselage drag share that is ingested into the propulsive device, while D Fus,not ing constitutes the outer part of the fuselage viscous flow field that is spilled around the nacelle, e.g. in case the intake duct height is smaller than the boundary layer thickness at the intake position. Substituting equation (7) into equation (6) produces: Fˆ ˆ D (8) N F DNac DFus, ing DFus, noting Since the individual contributors to Dˆ including D Fus and D Nac were not available from CFD results, these were parametrically determined based on the approach given by Raymer (2006). As a plausibility check, D Fus was compared to the corresponding CFD result obtained for the clean fuselage geometry which yielded good jsf

8 Intake Duct Height Boundary Layer Thickness 8 J. Bijewitz, A. Seitz and M. Hornung agreement (order of +1% deviation). As it is of interest to quantify the fuselage drag share that is actually ingested into the FF, a semi-empirical boundary layer method based on Seitz and Gologan (2014) was utilized. Here, numerically computed boundary layer results derived for a generic elliptical fuselage geometry (van Dyck, 2012) had been employed. In order to account for the contraction of the fuselage towards the FF intake as considered in the present configuration, the boundary layer thickness obtained from the semi-empirical model was calibrated with the respective value measured from the newly obtained CFD results. The approach for calculating the fuselage drag share that is not ingested is based on the decomposition of the incoming flow into different shares of fluid momentum. In Figure 6, a boundary layer profile is schematically shown for a case where the intake duct height is smaller than the boundary layer thickness. The momentum of the complete boundary layer (profile assumed to be developed until 99% V 0 ) is approximated through the product of area averaged mean intake velocity, V 1,m, and the mass flow that is captured in the corresponding layer, m 1 : I V 1 m m (9), 1 The shares corresponding to the momentum that is ingested into the propulsive device, I ing, as well as the momentum that is passing around the FF nacelle, I not ing, may be expressed in a similar fashion. The term I 0 represents the momentum equivalent to free stream condition (V 0 ). y Momentum share of complete boundary layer (I δ ) Momentum share not ingested (I not ing ) Free stream momentum extending to intake duct height (I 0,h ) u Ingested momentum share of boundary layer (I ing ) Momentum share equivalent to free stream condition (I 0 ) Figure 6: Scheme indicating momentum shares associated with the Fuselage Fan inflow As can be seen from Figure 6, the share that is not ingested is given by: I noting I I I I ) (10) 0 ( 0, h ing where I 0,h represents the share of the free stream momentum extending only to the intake duct height. From this, an ingestion factor, f β, may be defined measuring the share of fuselage drag ingested into the propulsor relative to the complete ingestion: 1.0, hintake f I noting (11) 1, h intake I0 I

9 Multi-Disciplinary Design Investigation of Propulsive Fuselage Aircraft Concepts 9 Hence, the fuselage drag that is ingested into the FF in equation (8) is given by D Fus,ing = f β D Fus. Knowing the thrust produced by the propulsor ( Fˆ ) as well as the individual contributors to the overall drag, equation (8) may be rearranged to determine the drag effect caused by nozzle jet shear flow: D jsf Fˆ D D D Fˆ (12) Nac Fus, ing Fus, noting N In gas turbine design and performance, losses occurring in the propulsion system streamtube upstream and downstream of the engine, are typically accounted for in the thrust calculation using an adequate streamtube factor (cf. Seitz, 2012). For highly integrated propulsion system arrangements as considered in the present context, this factor incorporates all effects occurring due to installation of the FF propulsion system and hence correlates the isolated gas turbine performance to the respective values obtained for the integrated arrangement. Thus, it is convenient to introduce an integration factor, f int, which not only captures the fuselage drag spilled around the FF nacelle, but also covers additional effects caused by the integration of the FF propulsion system. These include, e.g. friction losses associated with nozzle shear flow at the fuselage aft-cone. Accordingly, for input settings corresponding to the CFD setup, the net thrust obtained from the gas turbine model, F N,calc, is overestimating the value Fˆ N yielded from equation (8), since yet none of the integration effects named above are included in F N,calc. The integration factor quantifying these losses is declared as: FN, calc f int 1.0 Fˆ (13) N The sum of all losses occurring due to FF integration may be considered as an additional contribution to the overall drag balance: F Fˆ D Fˆ ( f 1) Fˆ f Fˆ (14) N, calc N int N int N int where D int denotes the drag due to the FF integration. As can be seen, this is consistent with the inequality constraint formulated in equation (13). Using the sample designs indicated in Figure 2, a nonlinear two-dimensional regression model was derived featuring sensitivity with the inlet duct height at the Aerodynamic Interface Plane (AIP, see Figure 5), h AIP, and the design fan pressure ratio, p t13 /p t2. The root mean squared error between the sample points and the corresponding model approximation was 0.6%, while the maximum occurred error was +0.9%. The resulting correlation is defined as: 1.0 p t f int haip haip, h AIP in [m] (15) pt 2 A contour plot showing lines of constant f int as a function of h AIP and FPR is presented in Figure 7. It was found that f int is primarily dependent on the size of the propulsor intake. Here, the dominating effect is rooted in the increase of the portion of the fuselage drag that is not ingested but spilled around the nacelle as the duct height is decreased. In addition, decreasing duct heights cause increased interaction between the flow field around the nacelle stagnation point and the FF inflow field. The dependence of f int on FPR is only modest. As can be seen from Figure 7, increasing FPR yields improvement of f int, which is considered to be caused by an increased suction effect of the FF, and thus reduced intake spillage drag for growing FPR. Based on the derived approximation models, iterative design laws were subsequently integrated into the GasTurb propulsion system sizing and performance suite. In order to evaluate the quality of the implemented approach, the results obtained for a representative set of parameters using the adapted gas turbine model (y GT ) were compared to the respective values available from the CFD computations (y CFD ). N

10 10 J. Bijewitz, A. Seitz and M. Hornung D 0 Data points investigated in CFD analysis Implemented regression model (f int = const.) D 3 Design Fan Pressure Ratio (FPR) [-] D 4 2 D D Study Settings: Operating Condition: FL350, M0.80, ISA Fuselage Geometry acc. to Isikveren et al. (2014) Validity: h AIP = [0.5m, 1.0m], FPR = [1.25, 1.55] Duct Height at AIP (h AIP ) [m] 1.1 D 2 Figure 7: Implemented approximation model for integration factor The obtained verification results are shown in Figure 8 in terms of relative deviation y GT - y CFD /y CFD. As important check parameters, specific fan power, intake mass flow, and, nozzle pressure ratio are displayed. As can be seen, the CFD results are matched with adequate accuracy by the integrated GasTurb model. The mean standard error is 1.2%, while maximum and minimum relative deviations are 2.2% and -1.6%, respectively. Relative Deviation ( y GT - y CFD )/y CFD ) [-] Spec. Fan Power (Mean rel. error = 7.67e-03) Intake Mass Flow (Mean rel. error = 1.82e-02) Nozzle Pressure Ratio (Mean rel. error = 4.30e-03) 10-4 D0 D1 D2 D3 D4 D5 Investigated Design Points Figure 8: Deviation of design point results obtained from adapted GasTurb model versus CFD results 4. Parametric Study Results for Fuselage Fan Power Plant This section describes the results obtained from parametric studies performed for the FF propulsion system and highlights important differences compared to a conventionally installed power plant Reference propulsion system The reference propulsion system is assumed to be installed conventionally, i.e. in the free stream. Basic cycle characteristics including TET and OPR were chosen according to the propulsion system model (unmixed flow geared turbofan) derived in a previous investigation (cf. Bijewitz et al., 2014). The modelling strategy refers to the approach described in Section 3.2. For the intake pressure ratio, core intake pressure ratio and fan polytropic

11 Multi-Disciplinary Design Investigation of Propulsive Fuselage Aircraft Concepts 11 efficiency, typical values corresponding to the advanced technology status were applied. A synopsis of essential parameters is provided in Table Discussion of study results In order to ensure a consistent comparison against the reference power plant, OPR and TET were kept identical. Further design parameters were selected as given in Table 1. In order to account for the expected degradation of fan performance emanating from the disturbed, non-uniform inflow condition, fan polytropic efficiency was assumed to be reduced by 2.0 percentage points (Seitz et al., 2014). Table 1: Settings for power plant design studies Parameter Unit Reference PPS Fuselage Fan PPS Turbine Entry Temperature (T 4) a [K] Overall Pressure Ratio (OPR) a [-] Axial Fan Inlet Mach number (M ax,2) [-] Fan Inlet Hub/Tip Ratio (HTR 2) [-] 0.29 Iterated acc. to intake duct height Intake Pressure Ratio (p t2/p t0) [-] Iterated acc. to Figure 4 Core Intake Pressure Ratio (p t22/p t21) [-] Fan polytropic efficiency (η Fan) [-] Base a Max. Climb at Top-of-Climb (FL350, M0.80, ISA) In Figure 9, the characteristics of a sizing study are presented for both the reference propulsion system and the FF power plant. Here, specific thrust was varied for different values of net thrust requirement. Overall efficiency (η ov ) was selected as a metric for comparison against the reference. The abscissa displays the intake area (A 2 ). As can be seen, the reference propulsion system exhibits no sensitivity between η ov and net thrust, thus reflecting the assumption of invariant component efficiencies as well as zero secondary power and customer bleed air extraction. As expected, decreasing levels of specific thrust yield improvements in propulsive efficiency, which are, however, incrementally compensated by decreasing levels of transmission efficiency, thereby causing the contours of constant overall efficiency to converge towards the lower end of F N /W 2 shown. In contrast to the reference engine, the FF power plant shows a significant dependency of overall efficiency with net thrust. This is primarily caused by the influence of the inlet duct height on the intake pressure ratio, whose characteristics are included in Figure 9 using dashed contour lines. As can be seen, the identified trend of p t2 /p t0 presented in Figure 4 is also reflected in the power plant model. Counterintuitively, for constant net thrust, reduced levels of specific thrust yield decreasing values of η ov within the range of parameters studied Spec. Thrust (F N /W 2 ) [m/s] Net Thrust [kn] Fan Diameter [m] Underwing Podded Geared Turbofan Overall Efficiency ( ov ) [-] Study Settings: Technology Level: EIS 2035 Operation Conditions: FL350, M0.80, ISA Reference PPS: p t2 /p t0 = Zero Off-take/Customer Bleed Scenario Net Thrust [kn] Spec. Thrust F N /W 2 [m/s] Fuselage Fan Propulsion System 45 Intake Pressure Ratio [-] Intake Area (A ) [m 2 ] Figure 9: Design study for Fuselage Fan and reference propulsion system

12 12 J. Bijewitz, A. Seitz and M. Hornung In order to deeper explore this behavior, a detailed analysis of the FF propulsion system was conducted and the characteristics are given in Figure 10, where contours of net thrust and specific thrust are presented as a function of design fan pressure ratio and the intake duct height at the fan plane. As a consequence of the above illustrated strong dependence of the characteristics with propulsor size, the improving intake pressure ratio yields decreasing fan pressure ratios for equal specific thrust levels. As can be seen, decreasing values of specific thrust cause the transmission efficiency to degrade. As an inherent characteristic of ducted propulsive devices, the impact of pressure losses in the transmission system scales inversely proportional to F N /W 2. While for the reference power plant the increase in propulsive efficiency still yields increasing propulsive device efficiencies (η pd = η tr η pr ) (which propagates to increased overall efficiencies) within the range of F N /W 2 values investigated, the increased stream tube losses of the FF power plant exceed the respective losses occurring for the free stream case, i.e. podded power plants. Hence, the transmission efficiency is penalized more severely, thereby counteracting the improvement of propulsive efficiency stronger than for the reference case. This behavior reflects the findings gained from a previously conducted PF investigation (Seitz and Gologan, 2014). Transmission Efficiency ( tr ) Design Fan Pressure Ratio (FPR) [-] Spec. Thrust (F N /W 2 ) [m/s] 75 Study Settings: Technology Level: EIS Operation Conditions: FL350, M0.80, ISA Zero Off-take/Customer Bleed Scenario Net Thrust [kn] Intake Duct Height at Fan Plane (h 2 ) [m] Figure 10: Characteristics of Fuselage Fan Propulsion System 5. Integrated Performance Assessment As a final step, the impact of the investigated propulsion system characteristics on the integrated performance of the PF aircraft was studied. Here, emphasis was placed on identifying potential differences relative to a previous investigation which had been based on pure semi-empirical boundary layer methods (Seitz et al., 2014; Isikveren et al., 2014). Therefore, surrogate models for convenient mapping of engine design and off-design behavior were derived for the FF power plant model developed in the present context using the approach described in (Seitz, 2012). For the conceptual sizing and performance evaluation of the PF aircraft, the aircraft conceptual design methodology discussed in (Seitz, 2012) was employed. Implemented extensions required for the handling of PF concepts are outlined by Seitz et al. (2014). Generally, in the present context, the aircraft performance evaluation was conducted for the three-engine aircraft configuration given in Figure 1 for an air transport task of design range 4800 nm accommodating 340 passengers. In particular, the FF geometric arrangement was retained, which had been identified as suitable for the PF concept (Seitz et al., 2014). The results were compared to the characteristics of an advanced twin-engine reference aircraft sized for identical requirements and technology status, which is discussed in detail in (Seitz et al., 2014). Maximum wing loadings were retained constant for similar low-speed performance and common wing spans of 65.0 m were applied to ensure similar airport compatibility. The characteristics of the podded power plants

13 Multi-Disciplinary Design Investigation of Propulsive Fuselage Aircraft Concepts 13 were adopted unchanged. A common core strategy for the under wing podded and the FF power plants was implemented, i.e. the thrust split ratio was iterated to yield identical LPT design power outputs for all engines installed on the PF aircraft. A comparative synopsis of important propulsion system and aircraft related parameters is provided in Table 2. The FF core intake pressure ratio resulted from a parametric pressure loss model derived for a generic s-shape geometry, which was based on basic 1D flow relations. Regarding the weight estimation of the FF power plant, a number of refinements were implemented yielding an improved level of detail relative to the values predicted in (Seitz et al., 2014). This includes e.g. the parametric mapping of the weight of the s-duct and the intake structural elements. As can be seen from Table 2, the ingested drag ratio is significantly increased relative to results obtained from the semi-empirical method. While in Seitz et al. (2014) additional drag occurring due to the presence of the FF power plant integration was included in the prediction of ingested drag using a conservatively assumed constant factor, in the present method, integration losses were directly considered as part of FF power plant sizing via the integration factor, f int, introduced in Section Hence, β obtained using the present method is increased and is propagated to an increased apparent lift-to-drag ratio, while f int contributes to the degraded overall thrust specific fuel consumption (TSFC) relative to the semiempirical results. Due to the increased FF power plant weight and resulting vehicular cascade effects, e.g. increased wing reference area, Operating Weight Empty (OWE) is penalized more strongly than calculated before. This translates into an increase in Maximum Take-off Weight (MTOW) of 1.2% relative to the reference aircraft. The substantially reduced aircraft apparent drag still outweighs the BLI-induced degradation of power plant performance as well as the weight increase primarily caused by the third aft-installed engine. As a consequence, a block fuel burn reduction of 9.4%, or 10.4% relative improvement in ESAR, is obtained over the advanced reference aircraft for the selected PF design. The deltas in block fuel and ESAR relative to the previously obtained results are below 1%. In summary, important aircraft characteristics including the fuel burn improvement proved to be confirmed against semi-empirical methods. Table 2: Comparative synopsis of integrated aircraft characteristics using semi-empirical and CFD-derived methods Parameter Unit Semi-empirical methods (Seitz et al., 2014) Present method Delta a [%] FF Inlet Duct Height [m] ±0.0 FF Diameter [m] ±0.0 FF Specific Thrust [m/s] FF Intake Pressure Ratio [-] FF Core Intake Pressure Ratio [-] Delta FF Design Fan Efficiency [-] ±0.0 FF Propulsion System Weight [kg] Total TSFC at typical cruise c [g/s/kn] Ingested Drag Ratio [-] Apparent Lift-to-Drag Ratio d [-] Wing Reference Area [m 2 ] Delta OWE e [%] b Delta MTOW e [%] b Delta Block Fuel Burn e [%] b Delta ESAR e [%] +9.8% b a Results obtained from present method compared to semi-empirical methods b Percentage points c FL350, M0.80, ISA d at C L = 0.55, M0.80 e Relative to advanced reference aircraft (cf. Seitz et al., 2014) 6. Conclusion and Future Work In this paper, an analytical procedure for the matching of numerically calculated characteristics of a Propulsive Fuselage layout to the respective propulsion system was discussed. Based on the results of an initial CFD-based design space exploration, important heuristics specific to a Fuselage Fan propulsion system were identified and integrated into a power plant sizing and performance model. Thereafter, parametric design studies of the Fuselage Fan propulsion system were conducted. The results were compared and contrasted to the parametric

14 14 J. Bijewitz, A. Seitz and M. Hornung design results of a conventionally installed advanced turbofan architecture. Finally, the implication of the newly derived design heuristics on the integrated performance of the Propulsive Fuselage aircraft concept was studied. Here, emphasis was placed on identifying potential differences relative to a previous investigation which had been based on pure semi-empirical boundary layer methods. As a result, the aircraft-level benefit originally predicted using semi-empirical methods could be confirmed using the CFD-derived propulsion system characteristics yielding a fuel burn benefit of -9.4% (or relative improvement in ESAR of +10.4%) over the advanced reference aircraft. Future work will focus on a detailed exploration of the design space feasible for a Propulsive Fuselage aircraft. While the present paper considered a certain split of net thrust between the aft-installed and the podded power plants, this particularly will include variation of the thrust split. Additionally, the sensitivity of the investigated Propulsive Fuselage concept to important parameters such as design flight speed or Fuselage Fan design efficiency will be analyzed. Uncertainties of models and technical assumptions immanent at this early stage of technology evaluation will be treated using non-deterministic methods. Moreover, future work should also explore alternative solutions for the integration of fuselage fan power supply. This includes novel ways of power transmission to the large fuselage propulsor, for example through (hybrid-) electric power train options. Acknowledgements The authors would like to thank Dr. Askin T. Isikveren for fruitful discussions and valuable advice. Special gratitude is conveyed to Richard Grenon and Jean-Luc Godard, ONERA, as well as Stefan Stückl, Airbus Group Innovations, for their CFD analysis and CAD generation effort, respectively, which was conducted within the DisPURSAL project. This research was performed within the FP7-L0 project DisPURSAL (Grant Agreement No. FP ), co-funded by the European Commission. References Advisory Council for Aviation Research and Innovation in Europe. (2012). Strategic Research and Innovation Agenda. Brussels Atinault, O., Carrier, G., Grenon, R., Verbecke, C., Viscat, P. (2013). Numerical and Experimental Aerodynamic Investigations of Boundary Layer Ingestion for Improving Propulsion Efficiency of Future Air Transport. Proceedings of the 31st AIAA Applied Aerodynamics Conference, San Diego, California, 2013 Bijewitz, J., Seitz, A., Hornung, M. (2014). Architectural Comparison of Advanced Ultra-High Bypass Ratio Turbofans for Medium to Long Range Application. Document ID , Deutscher Luft- und Raumfahrtkongress 2014, Augsburg, Germany, 2014 Bolonkin, A. (1999). A high efficiency fuselage propeller ( Fusefan ) for subsonic aircraft. Word Aviation Conference, San Francisco, 1999 European Commission. (2011). Flightpath 2050: Europe s Vision for Aviation, Report of the High Level Group on Aviation Research. Luxembourg Grieb, H. and Schubert, H. (Ed.) (2004). Projektierung von Turboflugtriebwerken, Birkhäuser Verlag, Basel-Boston-Berlin Isikveren, A.T. (2012). Distributed Propulsion and Ultra-high By-pass Rotor Study at Aircraft Level (DisPURSAL). FP7- AAT-2012-RTD-L0, Proposal No , European Commission Directorate General for Research and Innovation Isikveren, A.T., Seitz, A., Bijewitz, J., Hornung, M., Mirzoyan, A., Isyanov, A., Godard, J.-L., Stückl, S., van Toor, J. (2014). Recent Advances in Airframe-Propulsion Concepts with Distributed Propulsion, Proceedings of the 29th Congress of the International Council of the Aeronautical Sciences, St. Petersburg, Russia, September 7-12, 2014 Kaiser, S., Grenon, R., Bijewitz, J., Prendinger, A., Atinault, O., Isikveren, A., Hornung, M. (2014). Quasi-Analytical Aerodynamic Methods for Propulsive Fuselage Concepts. Proceedings of the 29th Congress of the International Council of the Aeronautical Sciences, St. Petersburg, Russia, September 7-12, 2014 Kim, H. D. (2010). Distributed propulsion vehicles. Proceedings of the 27th Congress of the International Council of the Aeronautical Sciences, Nice, France, September 19-24, 2010 Kurzke, J., Gasturb11, Compiled with Delphi 2007 on 27 January, 2010 Raymer, D. (2006). Aircraft Design: A Conceptual Approach, 4th edition, AIAA Education Series, American Institute of Aeronautics and Astronautics, Inc., New York, NY Reynolds, C. (1985). Advanced Propfan Engine Technology (APET) Single- and Counterrotation Gearbox / Pitch Change Mechanism, Final Report. Pratt & Whitney United Technologies Corporation, NASA CR , Vol. 1 & 2

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